Multi-piece assembly for a tubular composite body

ABSTRACT

Embodiments are directed to systems and methods for two or more cured composite assemblies that are bonded together to form a tubular composite structure, wherein each of the cured composite assemblies do not have a tubular shape. The tubular composite structure may form a spar for an aerodynamic component, for example. The two or more cured composite assemblies may comprise carbon or fiberglass composite materials or a combination of materials. Each of the cured composite assemblies may further comprise axial edges that are configured to be bonded to another of the cured composite assemblies, wherein the axial edges have a sloped shape. An adhesive agent may be applied on the axial edges for bonding two cured composite assemblies. Alternatively, or additionally, one or more fasteners may be used to attach the axial edges of at least two cured composite assemblies.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a divisional of, and claims benefit of, U.S. patentapplication Ser. No. 16/048,869, filed Jul. 30, 2018, which is herebyincorporated herein by reference in its entirety.

BACKGROUND

Composite assemblies are created by laying up an assembly of uncureddetails and material. This typically consists of laying dry fabriclayers (“plies”) by hand to create a laminate stack. Resin is thenapplied to the dry plies after layup is complete. Alternatively, “wet”composite plies that have resin built in may be used in the layup.Composite fabrication usually involves some form of mold tool to shapethe plies and resin. A mold tool is required to give the unformedresin/fiber combination its shape prior to and during cure. Once thelayup is complete, the composite is cured. The cure can be acceleratedby applying heat and pressure to the composite layup.

A composite assembly may be used as a structural member for an aircraftcomponent, for example. These structural members are often referred toas a “spar,” and they may extend the axial length of a structure toprovide support against loads applied on the structure. In the case ofan aerodynamic component, such as propellers, rotor blades, and wings,for example, the spar may support both the weight of the aerodynamiccomponent and any aerodynamic loads applied to the aerodynamiccomponent, such as lift and drag forces. The spar is the primarystructural member or backbone of many aircraft components. Due to thetubular geometry of typical spars, it can be challenging to produce aspar that fully forms to the desired shape without wrinkles or otherdefects that arise due to the inherent trapping condition exhibited bynon-symmetric shapes and woven composite materials.

In existing manufacturing processes, a spar may be formed using acomposite preform that is cured prior to assembly with the othercomponents of the structure, such as skin assemblies in the case ofcomposite blades. During this curing process, an inflatable bladder maybe disposed within the uncured spar and expanded to help compact thelayers of preformed composite material and remove any excess air bubblesor other voids included in the preform as the spar is cured at anelevated temperature within a precision mold. Once cured, the othercomponents or details of the composite assembly are assembled with thespar. For instance, in the case of a rotor blade, outer skins and aleading edge are assembled with the spar and then bonded in a subsequentcuring process.

The process of laying up a spar as one single structure requires a lotof manipulation which can lead to defects during the manufacturingprocess. For example, when the plies in the layers are oriented atvarious angles, such as off-axis plies that overlie unidirectional,full-span plies, the difference can cause wrinkling and bunching of thelayers during cure.

SUMMARY

Embodiments are directed to systems and methods for two or more curedcomposite assemblies that are bonded together to form a tubularcomposite structure, wherein each of the cured composite assemblies donot have a tubular shape. The tubular composite structure may form aspar for an aerodynamic component, for example. The two or more curedcomposite assemblies may comprise carbon or fiberglass compositematerials or a combination of materials. Each of the cured compositeassemblies may further comprise an axial edge that is configured to bebonded to another of the cured composite assemblies, wherein the axialedge has a sloped shape. An adhesive agent may be applied on the axialedge for bonding two cured composite assemblies. Alternatively, oradditionally, one or more fasteners may be used to attach the axialedges of at least two cured composite assemblies.

One or more composite plies may be used to cover a seam where the two ormore cured composite assemblies are bonded together. One or morecomposite plies may be wrapped around the two or more cured compositeassemblies after they have been bonded together. The one or morecomposite plies wrapped around the two or more cured compositeassemblies may be cured while the cured composite assemblies are bondedtogether.

An example method for manufacturing composite assemblies compriseslaying up plies on molds for two or more composite assemblies, whereineach of the cured composite assemblies do not have a tubular shape,curing the two or more composite assemblies, and bonding the two or morecured composite assemblies together to form a tubular compositestructure. The method may further comprise forming an axial edge havinga sloped shape on the composite assemblies and mating the sloped axialedges together when bonding the two or more cured composite assemblies.The method may further comprise bonding the two or more cured compositeassemblies together using an adhesive agent on an axial edge. The methodmay further comprise attaching at least two cured composite assembliestogether using fasteners.

The method may further comprise applying one or more composite pliesover a seam where the two or more cured composite assemblies are bondedtogether. The method may further comprise wrapping one or more compositeplies around the two or more cured composite assemblies after they havebeen bonded together. The one or more composite plies wrapped around thetwo or more cured composite assemblies may be cured while the curedcomposite assemblies are bonded together. The tubular compositestructure may form a spar for an aerodynamic component, for example. Thetwo or more cured composite assemblies may comprise carbon, fiberglass,or other composite materials or a combination of materials.

BRIEF DESCRIPTION OF THE DRAWINGS

Having thus described the invention in general terms, reference will nowbe made to the accompanying drawings, which are not necessarily drawn toscale, and wherein:

FIG. 1 illustrates an example aircraft that can be used with certainembodiments of the disclosure.

FIG. 2 is a perspective view of an exploded uncured composite assemblyfor use in one embodiment.

FIG. 3A illustrates two halves of a generally symmetrical tubularcomposite part.

FIG. 3B illustrates a final tubular part once the halves shown in FIG.3A have been bonded together.

FIG. 4A illustrates multiple component assemblies of an asymmetricaltubular composite part.

FIG. 4B illustrates a final tubular part once the component assembliesshown in FIG. 4A have been bonded together.

While the system of the present application is susceptible to variousmodifications and alternative forms, specific embodiments thereof havebeen shown by way of example in the drawings and are herein described indetail. It should be understood, however, that the description herein ofspecific embodiments is not intended to limit the system to theparticular forms disclosed, but on the contrary, the intention is tocover all modifications, equivalents, and alternatives falling withinthe spirit and scope of the present application as defined by theappended claims.

DETAILED DESCRIPTION

Illustrative embodiments of the system of the present application aredescribed below. In the interest of clarity, not all features of anactual implementation are described in this specification. It will ofcourse be appreciated that in the development of any such actualembodiment, numerous implementation-specific decisions must be made toachieve the developer's specific goals, such as compliance withsystem-related and business-related constraints, which will vary fromone implementation to another. Moreover, it will be appreciated thatsuch a development effort might be complex and time-consuming but wouldnevertheless be a routine undertaking for those of ordinary skill in theart having the benefit of this disclosure.

In the specification, reference may be made to the spatial relationshipsbetween various components and to the spatial orientation of variousaspects of components as the devices are depicted in the attacheddrawings. However, as will be recognized by those skilled in the artafter a complete reading of the present application, the devices,members, apparatuses, etc. described herein may be positioned in anydesired orientation. Thus, the use of terms to describe a spatialrelationship between various components or to describe the spatialorientation of aspects of such components should be understood todescribe a relative relationship between the components or a spatialorientation of aspects of such components, respectively, as the devicedescribed herein may be oriented in any desired direction.

Embodiments are directed toward providing a high-quality composite partusing a process that lowers the risk of manufacturing defects andreduces the manufacturing time. A tubular composite assembly may be laidup in pieces that are later combined, which provides both qualityimprovements and potential manufacture time reductions. This providesoverall cost savings and allows for faster production rates.

FIG. 1 . illustrates an aircraft 101. Certain embodiments of thedisclosure may be used with an aircraft, such as aircraft 101. However,aircraft 101 is used merely for illustration purposes. It will beunderstood that composite materials manufactured using the embodimentsdisclosed herein may be used with any aircraft, including fixed wing,rotorcraft, commercial, military, or civilian aircraft, or any othernon-aircraft structure requiring a hollow or tubular construction.Embodiments of the present disclosure are not limited to any particularsetting or application, and embodiments can be used with a rotor systemin any setting or application such as with other aircraft, vehicles, orequipment. Certain embodiments of the composite assemblies and methodsof forming such disclosed herein may be used for any applicationinvolving a composite, aerodynamically shaped object. For example, someembodiments of the composite assemblies disclosed herein may be used forthe rotors, propellers, wings, or control surfaces of an aircraft.

Aircraft 101 may include fuselage 102, landing gear 103, and wings 104.A propulsion system 105 is positioned on the ends of wings 104. Eachpropulsion system 105 includes an engine 106 and a proprotor 107 with aplurality of rotor blades 108. Engine 106 rotates proprotor 107 andblades 108. Proprotor 107 may include a control system for selectivelycontrolling the pitch of each blade 108 to control the direction,thrust, and lift of aircraft 101. Although FIG. 1 shows aircraft 101 ina helicopter mode wherein proprotors 107 are positioned substantiallyvertical to provide a lifting thrust. It will be understood that inother embodiments, aircraft 101 may operate in an airplane mode whereinproprotors 107 are positioned substantially horizontal to provide aforward thrust. Proprotors 107 may also move between the vertical andhorizontal positions during flight as aircraft 101 transitions between ahelicopter mode and an airplane mode. Wings 104 may provide lift toaircraft 101 in certain flight modes (e.g., during forward flight) inaddition to supporting propulsion systems 105. Control surfaces 109 onwing 104 and/or control surfaces 110 are used to adjust the attitude ofaircraft 101 around the pitch, roll, and yaw axes while in airplanemode. Control surfaces 109 and 110 may be, for example, ailerons, flaps,slats, spoilers, elevators, or rudders. Wings 104, rotor blades 108,and/or control surfaces 109, 110 may be composite assemblies eachcomprising a spar and a set of upper and lower skins that extend alongthe spar. In some embodiments, the composite assemblies may have anupper core, a lower core, and a septum support layer extending betweenthe upper and lower cores.

FIG. 2 is a perspective view of an exploded uncured composite assembly201. In one embodiment, assembly 201 may be used to form the main rotorblades 108 of aircraft 101, for example. In another embodiment, assembly201 may be used to form the wings 104 and/or control surfaces 109, 110of aircraft 101. Composite assembly 201 generally comprises a pluralityof details, such as a spar 202, a trailing-edge core 203, an upper skin204, a lower skin 205, a leading-edge sheath 206, and an abrasion strip207. The core and skin structures may be bonded or otherwise attached tothe spar 202 to create a desired airfoil profile. For example, the bladecomponents may be bonded together using layers of adhesive between eachinterface to form the final assembly 201.

Spar 202 itself may be a composite assembly, such as fabric layers orplies that are laid by hand to form a laminate stack and then curedusing a resin that is applied to the dry plies after layup is complete.Spar 202 may have a central cavity 208 to create a hollow structure toreduce weight. Spar 201 may comprise two or more layers of uncuredunidirectional laminate material. The plurality of unidirectional layersmay be stacked or layered at varying angular directions relative to oneanother to achieve the desired strength and flexibility desired for theparticular application. Each unidirectional layer is formed fromfiberglass or carbon fiber composite material. However, in otherembodiments the unidirectional layers may comprise other types ofcomposite materials. In existing assemblies, spar 201 is manufactured asa single unit.

In embodiments of the disclosure, the design and manufacture of tubularcomposite bodies, such as spar 201, may be broken into two or more partsin order to simplify the manufacturing process and to minimize defects.The tubular composite bodies may be any symmetric and nonsymmetrictubular shape or composite body of revolution in which the fullcircumference design is divided into multiple pieces. When manufacturedas a single composite tubular component having plies that are orientedat different angles in different layers, the difference in plyorientation can cause wrinkling and bunching (i.e., “finger-trappingeffect”) of the layers during cure.

In one embodiment, a multi-piece assembly for a complex compositetubular assembly or body of revolution is laid up in two or moreseparate pieces, which are individually cured prior to assembly of thefinal part. The benefits of such a manufacturing process include ease ofthe layup process, lack of a trapping condition exhibited in manytubular-shaped parts, time-consuming hot compactions are not needed, andcommon defects inherent to the legacy one-piece composite partmanufacturing process are reduced. The finger-trapping effect, whichcommonly results in wrinkles in tubular composite assemblies due to thecircumferential nature of the single-part layup, is non-existent inmulti-piece assemblies since the composite fibers are not locked inplace.

FIG. 3A illustrates two halves 301, 302 of a generally symmetricaltubular composite part. FIG. 3B illustrates the final tubular part 300once the halves 301, 302 have been bonded together. Tubular compositepart 300 is first laid up and cured in separate “C”-shaped halves 301,302. Once the curing is complete, then the two halves 301, 302 arebonded together to form the final tubular-shaped part 300.

The final tubular part may be divided into any number of pieces. FIG. 4Aillustrates three components 401, 402, 403 of an asymmetric tubularcomposite part. FIG. 4B illustrates the final tubular part 400 once theparts 401, 402, 403 have been bonded together. The shape and number ofthe component parts 401, 402, 403 are tailored depending upon thecomplexity of the geometry of the final part 400 and the requirements ofeach individual component part 401, 402, 403.

The component parts may be bonded together using adhesive and/orfasteners. The component parts may be constructed to enhance orotherwise support bonding together. For example, rather than havingsquared off edges, the edges 303 of component parts 301, 302 may have ashallow angle or draft that increase the overlapping area between thetwo parts in order to maximize the bonding surface area. Edges 303 arereferred to herein as axial edges because they are oriented generallyparallel to the axis of the spar. Depending upon the number of subpartsand the precured details, the seams or bond lines could be locatedanywhere around the circumference of the final composite assembly.

Composite plies may be laid over the seams and cured to protect or hidethe seam and/or to reinforce the bond between component parts. Inanother embodiment, torque-wrap plies may be laid up around (i.e., outerwrap) and/or laid up inside (i.e., inner wrap) the final assembly of thecomponent parts. The torque-wrap plies may be co-cured at the finalassembly of the component parts.

An additional advantage of laying up separate component partsindividually instead of laying up the entire tubular assembly is theability to select inner or outer molds for each component part. When atubular composite assembly is created as a single unit, the tool istypically used to form an inner surface on which the plies are laid up.However, when individual composite assembly components are created, eachpiece of the final tubular assembly can be formed using a tool thatshapes either the inner or outer surface of that component. Moreover,one or more composite assembly components may be laid up on an innermold tool and one or more other composite assembly components may belaid up on an outer mold tool. This allows for optimal tool selectionfor each component part. Each layer of plies may be formed fromfiberglass, carbon fiber, or other composite materials or a combinationof two or more materials.

In various embodiments, the plies used to create each of the compositeassembly components may be laid up over a male tool and/or laid upinside a female tool. Alternatively, different composite assemblycomponents for the same final tubular assembly may be laid up using bothmale and female tools. The selection of a tool for a composite assemblycomponent is not available for existing tubular composite parts, whichare typically laid up surrounding a male tool. The use of different moldtools in embodiments disclosed herein allows for optimized manufacturingof each composite assembly component.

Although the example illustrated in FIGS. 3A/B and 4A/B refer toconstruction of a spar, it will be understood that the disclosedcomposite manufacturing process can be used for any other tubular orconical aircraft components, such as a spindle, grip, cuff, and thelike.

The foregoing has outlined rather broadly the features and technicaladvantages of the present invention in order that the detaileddescription of the invention that follows may be better understood.Additional features and advantages of the invention will be describedhereinafter which form the subject of the claims of the invention. Itshould be appreciated that the conception and specific embodimentdisclosed may be readily utilized as a basis for modifying or designingother structures for carrying out the same purposes of the presentinvention. It should also be realized that such equivalent constructionsdo not depart from the invention as set forth in the appended claims.The novel features which are believed to be characteristic of theinvention, both as to its organization and method of operation, togetherwith further objects and advantages will be better understood from thefollowing description when considered in connection with theaccompanying figures. It is to be expressly understood, however, thateach of the figures is provided for the purpose of illustration anddescription only and is not intended as a definition of the limits ofthe present invention.

What is claimed is:
 1. A method for manufacturing composite assemblies,comprising: laying up plies on molds for two or more compositeassemblies, wherein each of the cured composite assemblies do not have atubular shape; curing the two or more composite assemblies; and bondingthe two or more cured composite assemblies together to form a tubularcomposite structure.
 2. The method of claim 1, further comprising:forming at least one axial edge having a sloped shape on the compositeassemblies; and mating the sloped axial edges together when bonding thetwo or more cured composite assemblies.
 3. The method of claim 1,further comprising: bonding the two or more cured composite assembliestogether using an adhesive agent on at least one axial edge.
 4. Themethod of claim 1, further comprising: attaching at least two curedcomposite assemblies together using fasteners.
 5. The method of claim 1,further comprising: applying one or more composite plies over a seamwhere the two or more cured composite assemblies are bonded together. 6.The method of claim 1, further comprising: wrapping one or morecomposite plies around the two or more cured composite assemblies afterthey have been bonded together.
 7. The method of claim 1, furthercomprising: wrapping one or more composite plies around the two or morecured composite assemblies; and curing the one or more composite plieswrapped around the two or more cured composite assemblies while thecured composite assemblies are bonded together.
 8. The method of claim1, wherein the molds for the two or more composite assemblies comprise afemale tool.
 9. The method of claim 1, wherein the molds for the two ormore composite assemblies comprise both female tools and male tools. 10.The method of claim 1, wherein the tubular composite structure forms aspar for an aerodynamic component.
 11. The method of claim 1, whereinthe two or more cured composite assemblies comprise one or more ofcarbon and fiberglass composite materials.
 12. A method formanufacturing a rotor blade for an aircraft, comprising: providing twoor more cured composite components, each composite component having anouter surface; bonding the two or more composite components together toform a composite structure in which the outer surface of each compositecomponent forms at least a portion of an overall outer surface of thecomposite structure, wherein each of the two or more compositecomponents do not have a tubular shape; laying up composite plies insidethe two or more composite components; wrapping composite plies aroundthe two or more cured composite components; and co-curing the compositeplies and an assembly of the two or more cured composite componentsusing an adhesive agent.
 13. The method of claim 12, wherein the two ormore cured composite components comprise at least one of carbon andfiberglass composite materials.
 14. The method of claim 12, wherein eachof the cured composite components comprise an aerodynamic componentconfigured to be bonded to another of the cured composite components,and wherein at least one axial edge has a sloped shape.
 15. The methodof claim 14, further comprising: applying an adhesive agent on the atleast one axial edge to bond two of the cured composite components. 16.The method of claim 12, further comprising: bonding at least one axialedge of the cured composite components to another of the cured compositecomponents; and attaching the axial edges of the at least two curedcomposite component with fasteners.
 17. The method of claim 12, furthercomprising: wrapping one or more composite plies around an assemblycomprising two or more cured composite components.
 18. The method ofclaim 12, wherein the composite structure forms a spar for the rotorblade.